The present invention relates to the general field of combustion chambers for an aviation or terrestrial turbomachine.
Typically, an aviation or terrestrial turbomachine comprises an assembly made up in particular of: an annular compression section for compressing the air that passes through the turbomachine; an annular combustion section located at the outlet from the compression section and in which the air coming from the compression section is mixed with fuel in order to be burnt therein; and an annular turbine section disposed at the outlet from the combustion section and having a rotor that is driven in rotation by the gas coming from the combustion section.
The compression section is in the form of a plurality of rotor wheel stages, each carrying blades that are located in an annular channel through which the turbomachine air passes and of section that decreases from upstream to downstream. The combustion section comprises a combustion chamber in the form of an annular channel in which the compressed air is mixed with fuel in order to be burnt therein. The turbine section is made up of a plurality of rotor wheel stages, each carrying blades that are located in an annular channel through which the combustion gas passes.
The flow of air through the above assembly generally takes place as follows: the compressed air coming from the last stage of the compression section possesses natural gyratory motion with an angle of inclination of the order of 35° to 45° relative to the longitudinal axis of the turbomachine, which angle of inclination varies as a function of the speed of the turbomachine (speed of rotation). At its inlet into the combustion section, this compressed air has its flow straightened to become parallel with the longitudinal axis of the turbomachine (i.e. the angle of inclination of the air relative to the longitudinal axis of the turbomachine is brought back to 0°) by means of air flow straightener vanes. The air in the combustion chamber is then mixed with the fuel so as to provide satisfactory combustion, and the gas produced by the combustion continues to flow generally along the longitudinal axis of the turbomachine so as to reach the turbine section. In the turbine section, the combustion gas is redirected by a nozzle so as to present gyratory motion with an angle of inclination greater than 70° relative to the longitudinal axis of the turbomachine. Such an angle of inclination is essential for producing the angle of attack needed to provide the mechanical force for imparting rotary drive to the rotor wheel of the first stage of the turbine section.
Such an angular distribution for the air passing through the turbomachine presents numerous drawbacks. The air that leaves the last stage of the compression section naturally presents an angle lying in the range 34° to 45° and its flow is successively straightened out (angle returned to 0°) on entering into the combustion chamber, and is then redirected to have an angle greater than 70° at its entry into the turbine section. Those successive changes to the angle of inclination of the air flow through the turbomachine require intense aerodynamic forces to be produced by the flow-straightener vanes of the compression section and by the nozzle of the turbine section, which aerodynamic forces are particularly harmful to the overall efficiency of the turbomachine.